It is known that a blade for an aircraft rotating aerofoil usually includes an aerodynamically profiled main part, extending, along the blade chord, between a leading edge and a trailing edge, and, along the span of the blade, between a footing part equipped with a blade root for its connection to a rotor hub rotating around a rotor rotation axis, and a blade tip, at its free end.
During rotor rotation, vortexes are generated at the blade tip by the pressure difference between the intrados and the extrados at each blade tip. In certain flight configurations, particularly in the low speed approach phase, these vortexes can interact with the blades following the blade emission of vortexes, and the pressure gradients on the blades, created by the passage of a vortex in the plane of the blades, generate an impulsive noise which can be very intense.
This impulsive noise source results from the fact that the leading edge of a blade enters simultaneously into contact, along practically all its span, with the vortexes generated by at least one previous blade.
It is known that such interactions can be produced just as easily on a forward moving blade as a backward moving blade.
From the side of the forward moving blade, the vortexes are emitted in the second quadrant (blade azimuths between 90.degree. and 180.degree., the azimuth 0.degree. corresponding by convention to the position of the rear blade) and the interactions take place in the first quadrant (azimuths between 0.degree. and 90.degree.). The sound emission has a directivity in the direction of helicopter forward movement, under the main rotor.
From the side of the backward moving blade, the vortexes are emitted in the third quadrant (azimuths between 180.degree. and 270.degree.) and the interactions take place in the fourth quadrant (azimuths between 270.degree. and 360.degree.). The noise produced is directed toward the rear of the helicopter, under the main rotor.
The blade-vortex interaction noise is very disadvantageous, for it is maximum for angles of descent which correspond to those adopted by helicopters in the approach phase during landings.
The main problem at the basis of the invention is to decrease the blade-vortex interaction noise in particular in a range of descent angles which covers angles for which the blade-vortex interaction noise emitted is maximum (angle of descent of about 6.degree.).
Different means have already been proposed in order to reduce the blade-vortex interaction noise, which depend on parameters of two main types, some linked to the vortex itself, and others to the geometry of the blade-vortex interaction. The different known means tend therefore either to modify the characteristics of the vortexes emitted, that is to say mainly their intensity and their viscous radius at the moment of interaction, or to modify the geometry of the interaction, mainly determined by the distance between the blade and vortex and the vertical and horizontal angles between the blade and the vortex line at the moment of interaction, these different parameters being determined mainly by the flight conditions, the power of the rotor and its rotational speed, the number of blades which it includes and their geometry (plane form, twisting, aerodynamic profile and blade tip).
Amongst the known means proposed for modifying the characteristics of the vortex at emission, distinction can be made between a first family of means aiming to decrease the eddying intensity of the vortex generated at the blade tip, and a second family of means aiming to create a second vortex, more internal than the blade tip vortex, in order to distribute the eddying intensity.
The known means of the first family are themselves sub-divided into passive and active means, the use of which modifies the aerodynamic circulation distribution on the blade span, because the intensity of the emitted vortex at the blade tip is directly connected to the span circulation gradient on the blade in this place, this gradient being all the higher as the maximum local load is situated near the blade tip.
A first known passive means consists in slimming down the tip or the blade end, so as to displace the maximum local circulation toward the inside of the rotor, and different blade end shapes slimmed down on the chord have been proposed.
A second known passive means consists in applying to the blade a twisting law leading to a weak blade tip circulation gradient, at the emission azimuths.
Other known passive means consist in adding a vertical aileron on the end profile of the blade tip in order to prevent or disrupt the rolling up of the vortex, or to add a spoiler on the leading edge at the blade end, in order to increase the viscous radius of the vortex.
The known active means are of ejecting an air flow at the blade tip and toward the trailing edge, in order to diffuse the vortex, or to reduce the vortex intensity with the help of control laws of appropriate pitch, such as multi-cyclic controls and individual blade controls, enabling controlling the pitch of the blades so that the lift of the blades at the emission azimuth of the interaction vortexes is as low as possible.
The second family of known means, aiming to distribute the vortex intensity, includes the addition of a small wing at the blade tip, so as to force the generation of two vortexes, one at the blade tip, and the other at the end of the small additional wing.
The known means for modifying the characteristics of the blade-vortex interaction geometry in order to decrease the noise are also subdivided into active and passive means.
The active means include multi-cyclic controls, already mentioned, controlling, for this application, the pitch of the blades in order that it is at a maximum on the side of the forward moving blade, between the interaction azimuths and the emission azimuths, in order to increase the load and the induced speed downward, and therefore to accelerate the convection of vortexes downward, so that the following blades do not enter, or enter the least possible, into collision with the emitted vortexes.
The passive means consist in modifying the geometry of the leading edge or of the line passing through the aerodynamic centres of the successive basic streamlined sections of the blade along the span, called quarter chord line (because the aerodynamic centres are each usually situated in the front quarter of the chord of the corresponding basic blade section). Indeed, the blade-vortex interaction noise is all the greater as the acoustic peaks emitted by the different blade sections reach an observation point in phase. Therefore, when the blade, at a given azimuth, is parallel to a vortex line, the different blade sections enter into collision with the vortex at the same instant. In order to break the parallelism between a blade and a vortex line, and thus put out of phase the acoustic sources, a blade has been proposed the external part of which is front cambered, whereas a more internal rear cambered part balances the blade.
The blade-vortex interaction noise can also be reduced by modification of the descent conditions and/or the helicopter approach trajectory, by increasing the angle of descent and/or the helicopter speed, so as to move away from the noisiest configuration.
This solution is however applicable with difficulty, because, for comfort and safety reasons, it is not conceivable to use angles of descent greater than 8.degree., whereas the optimum angle of descent of 6.degree., which corresponds to the certification angle of descent of helicopters, is that in the neighbourhood of which the maximum noise is emitted. [The acoustic nuisance of a rotor can also be reduced by reduction of the speed of rotation and/or the increase of the number of blades, but the gains obtained depend strongly on the actual flight conditions, during which fluctuations of descent slope, of wind and of speed impose relatively independent solutions for the aforementioned different parameters.
The problem at the basis of the invention is to decrease the sound emission on the one hand by an attenuation of the impulsive characteristics of the noise source which is more pronounced than with the means known to that end, and additionally independent of the angle of descent, by reducing the blade-vortex interaction noise by modification of the interaction geometry, and by breaking the parallelism, particularly horizontal, between the vortex line and the leading edge of the blade.
Another aim of the invention is to reduce the noise of a rotating blade by also modifying the end vortex which the blade generates and/or by modifying the blade-vortex distance, in particular by breaking the vertical parallelism between the vortex lines and the blade.
The invention has the further aim of obtaining a reduction of the blade-vortex interaction noise without penalty to the aerodynamic performances of the blade in the whole of the flight domain.